EP2998519A1 - Turbinenmotordiffusoranordnung mit luftkanalmischer - Google Patents

Turbinenmotordiffusoranordnung mit luftkanalmischer Download PDF

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Publication number
EP2998519A1
EP2998519A1 EP15185959.2A EP15185959A EP2998519A1 EP 2998519 A1 EP2998519 A1 EP 2998519A1 EP 15185959 A EP15185959 A EP 15185959A EP 2998519 A1 EP2998519 A1 EP 2998519A1
Authority
EP
European Patent Office
Prior art keywords
mixing chamber
airflow
assembly
mixing
mixer
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP15185959.2A
Other languages
English (en)
French (fr)
Other versions
EP2998519B1 (de
Inventor
Frank J. Cunha
Stanislav KOSTKA
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
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Filing date
Publication date
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Publication of EP2998519A1 publication Critical patent/EP2998519A1/de
Application granted granted Critical
Publication of EP2998519B1 publication Critical patent/EP2998519B1/de
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Anticipated expiration legal-status Critical

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
    • F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16—Control of working fluid flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/005—Combined with pressure or heat exchangers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration

Definitions

  • This disclosure relates generally to a turbine engine and, more particularly, to a diffuser assembly for a turbine engine.
  • a typical turbine engine includes a combustor between a compressor and a turbine.
  • the combustor may be arranged within a diffuser module, which functions in general as an air distribution intermediary between the compressor and the combustor.
  • some of the compressed air within the diffuser module may be diverted for cooling other components of the turbine engine.
  • Various arrangements are known in the art for diverting this air; however, each of these arrangements has its own disadvantageous associated therewith.
  • a diffuser assembly for a turbine engine.
  • This diffuser assembly includes a diffuser module with a combustor plenum and a mixing chamber.
  • the diffuser module is configured to receive first and second airflows into the mixing chamber and direct a mixed airflow out of the mixing chamber.
  • the diffuser module includes a mixer configured to mix the first and the second airflows together within the mixing chamber to provide the mixed airflow.
  • an assembly for a turbine engine.
  • This turbine engine assembly includes a diffuser module with an inlet passage, a mixing chamber and first and second inlets leading into the mixing chamber.
  • the diffuser module includes a hollow strut and a mixer.
  • the hollow strut extends across the inlet passage and is fluidly coupled with the mixing chamber.
  • the mixer is configured to provide a flow obstruction between the first inlet and the hollow strut.
  • a method is provided involving a diffuser module of a turbine engine. This method includes receiving first and second airflows in a mixing chamber of the diffuser module. The first and the second airflows are mixed together within the mixing chamber using a mixer to provide a mixed airflow, which is provided to another component of the turbine engine.
  • the first airflow may be a conditioned airflow.
  • the second airflow may have a higher pressure than the first airflow.
  • the second airflow may also be conditioned.
  • a combustor may be included within a combustor plenum of the diffuser module.
  • the inlet passage may lead to the combustor plenum.
  • a hollow strut may be included that extends across an inlet passage of the diffuser module which inlet passage leads to the combustor plenum.
  • the hollow strut may be configured to receive at least a portion of the mixed airflow from the mixing chamber.
  • An inlet may be included through which at least a portion of the first airflow is directed into the mixing chamber.
  • the inlet may be substantially circumferentially aligned (or offset) with the hollow strut.
  • An inlet may be included through which at least a portion of the first airflow is directed into the mixing chamber.
  • the mixer may be configured to obstruct a direct path between the inlet and the hollow strut.
  • the mixer may be configured to diffuse the first airflow into the mixing chamber.
  • An inlet may be included through which at least a portion of the first airflow is directed into the mixing chamber.
  • the mixer may include or be configured as a mixing device located at the inlet.
  • the mixing device may be mated with the inlet.
  • the mixer may include or be configured as a mixing device with a cupped body and a plurality of perforations through the body.
  • Each of the perforations may have a substantially uniform configuration; e.g., size, shape, etc.
  • one of the perforations may have a different configuration than another one of the perforations.
  • the mixer may include or be configured as a plurality of discrete mixing devices. These mixing devices are configured to cumulatively mix the first and the second airflows together within the mixing chamber to provide the mixed airflow.
  • Each of the mixing devices may have a substantially uniform configuration. Alternatively, one of the mixing devices may have a different configuration than another one of the mixing devices.
  • a heat exchanger may be included and configured to provide the first airflow.
  • the heat exchanger may be fluidly coupled between the combustor plenum and the mixing chamber.
  • the diffuser module may be configured to direct the second airflow from the combustor plenum into the mixing chamber; e.g., without passing through a heat exchanger.
  • FIG. 1 is a side cutaway illustration of a geared turbine engine 10, which is configured as a turbofan engine.
  • the turbine engine 10 extends along an axial centerline 12 between an upstream airflow inlet 14 and a downstream airflow exhaust 16.
  • the turbine engine 10 includes a fan section 18, a compressor section 19, a combustor section 20 and a turbine section 21.
  • the compressor section 19 includes a low pressure compressor (LPC) section 19A and a high pressure compressor (HPC) section 19B.
  • the turbine section 21 includes a high pressure turbine (HPT) section 21A and a low pressure turbine (LPT) section 21B.
  • the engine sections 18-21 are arranged sequentially along the centerline 12 within an engine housing.
  • Each ofthe engine sections 18, 19A, 19B, 21A and 21B includes a respective rotor 22-26.
  • Each of these rotors 22-26 includes a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks.
  • the rotor blades may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s).
  • the fan rotor 22 is connected to a gear train 28, for example, through a fan shaft 30.
  • the gear train 28 and the LPC rotor 23 are connected to and driven by the LPT rotor 26 through a low speed shaft 31.
  • the HPC rotor 24 is connected to and driven by the HPT rotor 25 through a high speed shaft 32.
  • the shafts 30-32 are rotatably supported by a plurality of bearings 34; e.g., rolling element bearings and/or thrust bearings.
  • Each of these bearings 34 is connected to the engine housing by at least one stationary structure such as, for example, an annular support strut.
  • the air within the core gas path 36 may be referred to as "core air”.
  • the air within the bypass gas path 38 may be referred to as "bypass air”.
  • the core air is directed through the engine sections 19-21, and exits the turbine engine 10 through the airflow exhaust 16 to provide forward engine thrust.
  • fuel is injected into a combustion chamber 40 and mixed with the core air. This fuel-core air mixture is ignited to power the turbine section 21.
  • the bypass air is directed through the bypass gas path 38 and out of the turbine engine 10 through a bypass nozzle 42 to provide additional forward engine thrust.
  • at least some of the bypass air may be directed out of the turbine engine 10 through a thrust reverser to provide reverse engine thrust.
  • FIG. 2 illustrates an assembly 44 of the turbine engine 10.
  • This turbine engine assembly 44 includes a combustor 46 arranged within a combustor plenum 48 of a diffuser module 50.
  • This combustor plenum 48 receives compressed core air from the HPC section 19B through an inlet passage 52 of the diffuser module 50.
  • the combustor plenum 48 provides the received core air to the combustor 46 as well as to other components as described below in further detail.
  • the turbine engine assembly 44 also includes one or more fuel injector assemblies 54 arranged circumferentially around the centerline 12. Each of these fuel injector assemblies 54 includes a fuel injector 56 which may be mated with a swirler 58.
  • the fuel injector 56 injects the fuel into the combustion chamber 40.
  • the swirler 58 directs some of the core air from the combustor plenum 48 into the combustion chamber 40 in a manner that facilitates mixing the core air with the injected fuel.
  • Quench apertures (not shown) in inner and/or outer walls 60 and 62 of the combustor 46 may direct additional core air into the combustion chamber 40 for combustion. Additional core air may also or alternatively be directed (e.g., effused) into the combustion chamber 40 through cooling apertures (not shown) in the inner and the outer walls 60 and 62.
  • the combustor 46 may be configured as an annular floating wall combustor.
  • the combustor 46 of FIG. 2 for example, includes an annular combustor bulkhead 64, the tubular combustor inner wall 60, and the tubular combustor outer wall 62.
  • the bulkhead 64 extends radially between and is connected to the inner wall 60 and the outer wall 62.
  • Each wall 60, 62 extends axially along the centerline 12 from the bulkhead 64 towards the HPT section 21A, thereby defining the combustion chamber 40.
  • Each combustor component 60, 62, 64 may be a multi-walled structure that includes, for example, a heat shield attached to a shell.
  • One or more of these components 60, 62, 64 may alternatively each be a single walled structure; e.g., comprise a single liner.
  • FIG. 3 illustrates a diffuser assembly 66 configured with the turbine engine assembly 44.
  • This diffuser assembly 66 includes the diffuser module 50 configured with an airflow conditioning system 68.
  • the diffuser module 50 may include one or a plurality of various turbine engine components. Examples of such turbine engine components include, but not limited to, cases, shrouds, struts, vanes, seals, ducts, etc.
  • the diffuser module 50 and its various components may be configured as an annular body that extends along the centerline 12 to and/or about the combustor 46.
  • the diffuser module 50 forms the combustor plenum 48, the inlet passage 52 (see FIG. 2 ) and a mixing chamber 70.
  • the combustor plenum 48 of FIG. 2 is defined radially between an inner case 72 and an outer case 74.
  • the inlet passage 52 is defined by a pre-diffuser assembly 76. More particularly, the inlet passage 52 is defined radially between inner and outer shrouds 78 and 80 of the pre-diffuser assembly 76.
  • the pre-diffuser assembly 76 also includes one or more hollow struts 82 (e.g., hollow structural guide vanes), which extend radially across and circumferentially subdivide at least a portion of the inlet passage 52 as illustrated in FIG. 4 .
  • the mixing chamber 70 is defined radially between the pre-diffuser assembly 76 and the outer case 74.
  • the mixing chamber 70 and the combustor plenum 48 are generally fluidly separated by at least one annular wall 84 (e.g., a seal), which extends between the outer case 74 and the outer shroud 80 (see FIG. 2 ).
  • the wall 84 of FIG. 3 may include one or more axially extending orifices 86 which provide a leakage flowpath between the combustor plenum 48 and the mixing chamber 70; e.g., an inlet into the mixing chamber 70.
  • the wall 84 may include a plurality of arcuate sections and one or more circumferential gaps between these sections may be configured to provide the leakage flowpath. Still alternatively, the wall 84 may be omitted, for example, where an annular lip is arranged between the combustor plenum 48 and the mixing chamber 70.
  • Each of the hollow struts 82 forms an inner bore 88 (e.g., outlet passage) that fluidly couples the mixing chamber 70 with at least one flowpath 90 and/or manifold.
  • This flowpath 90 and/or manifold fluidly couples the inner bores 88 with one or more other components 92 of the turbine engine 10; e.g., one or more of the rotor 22-26 components.
  • the airflow conditioning system 68 may include a heat exchanger 94.
  • This heat exchanger 94 may include one or more heat exchanger units connected serially and/or in parallel.
  • An inlet of the heat exchanger 94 is fluidly coupled with the combustor plenum 48 through a flowpath 96 (e.g., a duct or hose) and/or manifold.
  • An outlet of the heat exchanger 94 is fluidly coupled with the mixing chamber 70 through a flowpath 98 (e.g., a duct or hose) and/or manifold.
  • the heat exchanger 94 is configured to cool (or heat) or otherwise condition an airflow received from the combustor plenum 48 and provide the conditioned airflow to the mixing chamber 70.
  • the flowpath 98 is fluidly coupled with the mixing chamber 70 through one or more inlets 100.
  • each of these inlets 100 may be substantially aligned, circumferentially and/or axially, with a respective one of the inner bores 88 and/or a respective one of the inlets 86.
  • one or more of the inlets 100 may be offset, circumferentially and/or axially, from the inner bores 88 and/or the inlets 86.
  • each of the inlets 100 may be positioned circumferentially between a respective adjacent pair of the inner bores 88.
  • the diffuser module 50 further includes a mixer 102.
  • This mixer 102 is configured to mix airflows entering the mixing chamber 70 from the inlets 86 and 100 together to provide a mixed airflow.
  • the mixer 102 may be arranged to obstruct a direct path between one or more of the inlets 100 and each corresponding inner bore 88. Air entering the mixing chamber 70 through the respective inlets 100 therefore must move circumferentially and/or axially within the mixing chamber 70, which movement promotes mixing with other air within the mixing chamber 70; e.g., air entering through the inlets 86.
  • the mixer 102 may also or alternatively be configured to diffuse air entering the mixing chamber 70 through the respective inlets 100, which again promotes mixing of the diffused air with other air within the mixing chamber 70.
  • the mixer 102 may include one or more discrete mixing elements 104. Each of these mixing elements 104 may be aligned between a respective one of the inlets 100 and a respective one of the inner bores 88. Each mixing element 104, for example, may be located at (e.g., on, in, adjacent or proximate) and mated with a respective one ofthe inlets 100.
  • each mixing element 104 may be configured with a perforated cupped body (e.g., as a sieve).
  • the geometry of the cupped body paired with the placement ofthe perforations 106 may be tailored to redirect certain portions of incoming air in various directions. In this manner, the mixing element 104 may divert the path ofthe incoming air away from the inner bores 88 as well as diffuse the incoming air to further promote mixing.
  • the perforations 106 through the cupped body may have a substantially uniform configuration; e.g., the same shape and size. Alternatively, one or a set ofthe perforations 106 may have a different configuration (e.g., shape and/or size) than another one or set of the perforations 106.
  • the mixing elements 104 may have a substantially uniform configuration. Alternatively, one or a set of the mixing elements 104 may have a different configuration than another one or set of the mixing element 104. Furthermore, in other embodiments, the mixing elements 104 may only be configured with a subset of the inlets 100. In some embodiments, one or more of the mixing elements 100 may also or alternatively be configured with one or more of the respective inlets 86.
  • compressed core air flows from the combustor plenum 48 into the heat exchanger 94 through the flowpath 96.
  • the heat exchanger 94 cools (or heats) and/or otherwise conditions the received core air.
  • the conditioned core air flows through the flowpath 98 and then the one or more inlets 100 and into the mixing chamber 70 as a conditioned airflow.
  • Pressure of the conditioned airflow is typically less than that of the compressed core air within the combustor plenum 48 due to a pressure drop across the heat exchanger 94 or another supply plenum. Therefore, to increase air pressure within the mixing chamber 70 and ensure proper airflow to the turbine engine components 92, some compressed core air is allowed to flow substantially directly from the combustor plenum 48 into the mixing chamber 70 through the one or more inlets 86 as a boost (e.g., high pressure) airflow. Temperature of this boost airflow, however, is typically much higher than that of the conditioned airflow. The mixer 102 therefore is provided to promote thorough mixing of the airflows together to provide a mixed airflow with a relatively even, but low temperature.
  • a boost e.g., high pressure
  • the conditioned airflow may have the tendency to flow directly from inlets 700 into hollow struts 702 without much if at any mixing within chamber 74 as illustrated in FIG. 7 .
  • This can lead to relatively high temperature differentials within the chamber 704 as well as across components (e.g., struts 702) of the diffuser, which can lead to relatively high thermal stresses and premature component failure.
  • upstream is used to orientate the components of the systems described above relative to the turbine engine 10 and its centerline 12.
  • downstream is used to orientate the components of the systems described above relative to the turbine engine 10 and its centerline 12.
  • inner is used to orientate the components of the systems described above relative to the turbine engine 10 and its centerline 12.
  • outer is used to orientate the components of the systems described above relative to the turbine engine 10 and its centerline 12.
  • a person of skill in the art will recognize, however, one or more of these components may be utilized in other orientations than those described above.
  • the present invention therefore is not limited to any particular spatial orientations.
  • the systems may be included in various turbine engines other than the one described above.
  • the systems may be included in a geared turbine engine where a gear train connects one or more shafts to one or more rotors in a fan section, a compressor section and/or any other engine section.
  • the systems may be included in a turbine engine configured without a gear train.
  • the systems may be included in a geared or non-geared turbine engine configured with a single spool, with two spools (e.g., see FIG. 1 ), or with more than two spools.
  • the turbine engine may be configured as a turbofan engine, a turbojet engine, a propfan engine, or any other type of turbine engine.
  • the present invention therefore is not limited to any particular types or configurations of turbine engines.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP15185959.2A 2014-09-19 2015-09-18 Turbinenmotordiffusoranordnung mit luftkanalmischer Active EP2998519B1 (de)

Applications Claiming Priority (1)

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US201462052825P 2014-09-19 2014-09-19

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EP2998519A1 true EP2998519A1 (de) 2016-03-23
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Cited By (1)

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EP3404209A1 (de) * 2017-05-18 2018-11-21 United Technologies Corporation Turbinenkühlanordnung

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US11732892B2 (en) * 2013-08-14 2023-08-22 General Electric Company Gas turbomachine diffuser assembly with radial flow splitters
US11236677B2 (en) * 2019-05-15 2022-02-01 Raytheon Technologies Corporation Diffuser case support structure

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WO2014052737A1 (en) * 2012-09-28 2014-04-03 United Technologies Corporation Inner diffuser case struts for a combustor of a gas turbine engine
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Cited By (2)

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Publication number Priority date Publication date Assignee Title
EP3404209A1 (de) * 2017-05-18 2018-11-21 United Technologies Corporation Turbinenkühlanordnung
US11268444B2 (en) 2017-05-18 2022-03-08 Raytheon Technologies Corporation Turbine cooling arrangement

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US20160084502A1 (en) 2016-03-24
US10378773B2 (en) 2019-08-13
EP2998519B1 (de) 2021-04-28

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